The present invention relates to space vehicles such as satellites and probes carrying electronic or optical payloads. Such vehicles are typically provided with means for generating, storing, and distributing electricity to power all of the on-board electronics. Depending on the electrical efficiency specific to each piece of electronic equipment, the equipment dissipates as heat a fraction of the electrical energy it consumes in operation. This heat must be distributed and channeled by conduction and must then be dumped from the space vehicle into the vacuum of space by radiation. The invention relates to heat radiators for performing this function.
Various configurations of fixed or deployable radiators are known that are of dimensions which vary with the size and the electrical power of satellites. A present trend in telecommunications satellites is to have ever increasing dimensions and also to have power generators of ever increasing power, thus giving rise to a need for increased radiating areas on board satellites.
A radiator is more effective if it is not illuminated by the sun; thus, for geostationary satellites, the fixed radiators are generally installed preferably on those panels of the satellite structure which are exposed to the north and to the south once the satellite is orbiting in its operating position and the body of the satellite is oriented so as to point its antennas to the surface of the earth.
Deployable radiators are used when the fixed areas inherent to the structure of the satellite are not sufficient to dump all of the heat that is generated on board. Like other deployable structures such as solar generators, antenna reflectors, etc., deployable radiators are folded against the main structure of the satellite so as to lie within the volume that is available under the nosecone of the launcher. After launch, the satellite separates from the launcher once the nosecone has been ejected from the launcher, and the various deployable structures are deployed as the need arises in the mission by deployment mechanisms that are specifically adapted to each case.
When launching a geostationary satellite, the launch rocket typically puts the satellite onto an injection orbit known as a geostationary transfer orbit (GTO).
Thereafter, propulsion means on board a geostationary satellite enable it to be transferred from the injection orbit to its final orbital position on the geostationary arc known as the geostationary earth orbit (GEO).
At present, these propulsion means are generally constituted by one or more chemical-burning engines integrated in the satellite.
In a future generation of satellites, in order to increase the capacity and the lifetime of the satellite, it is envisaged to generalize the use of electric thrusters for putting satellites onto station and for keeping them on station. Such a thuster is powered by energy that is renewable, i.e. electricity picked up by means of the solar generators, thus making it possible to limit the drawbacks associated with on-board chemical fuels, such as their mass and the space required to store them.
A new generation of electrically-propelled satellites is thus being designed and made. For the sequencing involved in putting the satellite onto station, i.e. transfer from its injection orbit to its final orbital position on the geostationary orbit, various options are possible.
When using electrical propulsion to perform all or part of this transfer, it is necessary to collect solar energy so as to be able to convert it into electrical energy in order to power the electric thrusters. This can be done by deploying the solar generators of the satellite in part or in full. Or else, in other scenarios, the satellite can perform its transfer between its injection orbit and the geostationary orbit or a portion of this transfer using energy that was already on board at launch, whether in chemical, electrochemical (battery), or even nuclear form.
In all cases, the energy available for transfer is limited by the energy collection or storage means which must be dimensioned as sparingly as possible.
Thus, in all cases, since the available energy is only just sufficient, it must not be wasted insofar as that is possible during the transfer stage.
Document D1=EP 0 780 304 A1 discloses deployable radiators, each having two essentially parallel surfaces, comprising a radiating first face and an opposite second face that is insulating, so that during transfer from GTO to GEO the deployable radiators are in the stored position, pressed against the structure of the satellite, with their radiating faces facing towards the structure of the satellite and their insulating faces facing towards the vacuum of space in order to reduce the onboard heating requirements during this stage, thereby economizing on-board energy. That system is illustrated in FIG. 1 and is described in greater detail below.
Other deployable radiator systems are known in the prior art, e.g. from document D2=WO 99/19212, and the teaching thereof is described below with reference to FIG. 2. In that case the deployable radiators have two radiating faces.
The deployable radiators of document D1 are very effective for thermal insulation during the transfer stage towards the geostationary orbit from the injection orbit. In contrast, once the satellite is on-station in the geostationary orbit, their effectiveness as heat radiators is compromised by the fact that approximately half of the available surface area is insulating and not radiating.
The deployable radiators of document D2 are radiating over both faces in full, but they provide no thermal insulation during the transfer stage.
The deployable radiators of the invention enable the drawbacks of the prior art to be mitigated. To this end, the invention provides a deployable radiator for a space vehicle, the radiator having a stored position prior to deployment and an operational deployed position after deployment; said radiator having two main faces comprising a first face which is stored towards the vehicle and an opposite, second face which faces towards space in the stored position; said first face being radiating over at least the major portion of its area; at least a portion of said second face being thermally insulating; wherein a large portion of said second face is thermally radiating, said portion being substantially hidden prior to deployment of said radiator by other elements of said space vehicle that act as a screen during launch and during at least a portion of the transfer of said vehicle into the geostationary orbit.
According to an advantageous characteristic, said deployable radiator is connected to the structure of the vehicle by a deployment mechanism having a hinge with a hinge axis, wherein said axis is positioned at a non-zero angle relative to the main axes of the structure of said vehicle. In a particular embodiment, said hinge is fixed on a portion of said space vehicle referred to as a xe2x80x9ccommunications modulexe2x80x9d, which module is dedicated mainly to the on-board payload equipment. In an alternative embodiment, said hinge is fixed on a portion of said space vehicle referred to as a xe2x80x9cservice modulexe2x80x9d, which module is dedicated mainly to the on-board equipment of the platform serving to support proper operation of the payload equipment.
In a preferred embodiment, when said radiator is in the established deployed position, it possesses at least one degree of freedom about at least one axis. In another embodiment, when said radiator is in the established deployed position, it possesses at least two degrees of freedom about at least two axes. In a particular embodiment, the main axis of said radiator in the deployed position is not parallel to a main axis of said vehicle. In another particular embodiment, the main plane of said radiator in the deployed position is not parallel to a main plane of said vehicle.